The performance of a gas/combustion turbine may be improved by increasing the temperature of the combustion process. As the firing temperature increases, the requirement for cooling of the turbine parts exposed to the combustion gases increases.
It is known to bleed a portion of the compressed air produced in the compressor section of a gas turbine for use as a cooling medium in the turbine portion of the engine. The compressed air may be injected into the flow of combustion gas to provide an insulating film along the turbine surfaces, or it may be passed through internal cooling passages formed in the hot turbine parts in a closed cooling system. After being heated in the internal cooling passages of a closed system, the heated compressed air may be returned to a lower pressure portion of the compressor, or alternatively, to the inlet of the combustor. By providing heated air to the combustor the overall efficiency of the gas turbine engine system may be improved. One example of such a prior art device is illustrated in U.S. Pat. No. 5,782,076 issued to Huber et al. on Jul. 21, 1998, incorporated by reference herein.
One drawback of the prior art device of Huber is that the return of the hot compressed air to the combustor may cause fluid flow disturbances that complicate the design of the combustor system. Another drawback of the aforementioned configuration is that it is limited to compressed air as the cooling medium. What is needed is an apparatus and method for cooling a combustion turbine wherein the waste heat from the turbine portion is returned to the combustion air without generating a flow disturbance. It is also desired to provide a method and apparatus for cooling hot turbine parts in a manner that provides an efficient transfer of heat away from the turbine parts while returning the waste heat to the combustion process.